1. Field of the Invention
This invention relates to deployable panel structures for spacecraft and similar applications in which high ratios of area to weight and deployed area to collapsed volume are desirable. The invention will be disclosed in connection with a deployable and retractable solar array of the type typically used to generate power aboard spacecraft. Specifically, this invention is for an improved structural design for large area solar arrays of ultra light weight which may be deployed and refolded in a space environment.
2. Prior Art
Deployable solar arrays utilizing solar cells mounted upon a foldable structure are well known in the art and widely used for the generation of power aboard spacecraft. These deployable solar arrays have typically been constructed of honeycomb structures or other light, rigid panels upon which the solar cells were mounted, the panels being hinged to one another on hinge lines perpendicular to the direction of deployment. In one such concept, rigid rectangular solar panels are arranged face to face in accordian fashion when in the collapsed position. Deployment is effected by means of an extensible boom which is extended from a central support, the outer end of the boom being attached to the outer panel of the solar array such that the solar panel is deployed to an essentially planar configuration by the extension of the boom.
As the requirements for more power for satellites have developed there has been increased pressure to develop advanced concepts which will reduce the weight and stowed envelope size of solar arrays in order to provide a greater ratio of power generated in orbit to folded size and weight of the array. These requirements led to attempts to reduce or eliminate the rigid panel structures, but these reductions have caused serious difficulty in control of panel geometry and deployment kinematics. In particular, as the rigidity of panels is reduced, it becomes progressively more difficult to prevent bending of the panels between hinges, especially if an attempt to fold the panels back to the predeployment configuration is made. This problem is primarily due to the fact that the normal means of deployment is to attach the outer end of the outermost panel to the outer end of an extensible boom, the inner end of which is mounted on the central support of the solar array. During deployment the extensible boom outer end travels away from the central support, unfolding the solar panel as it extends. Intermediate dynamics are relatively unimportant as the panel when deployed is stretched tight by the extended boom.
The advent of the space shuttle and its ability to recover spacecraft from orbit has created a requirement that spacecraft be recoverable and restowable in the space shuttle in order to be returned to earth for refurbishment. Furthermore, the desire to conduct a number of experiments housed in the space shuttle required that the shuttle itself be able to deploy and recover a solar panel capable of producing large amounts of electrical power.
Refolding the panels of such a solar array, however, present serious difficulties, even in a zero gravity environment. This is due to the fact that intermediate panels between the end panels, if uncontrolled by additional mechanism, may pivot away from the boom, thereby defeating the restowing attempt. One means of preventing such an undesired kinematic result is to preload the hinge means to the stowed position so that the retracting of the boom is not the primary stowing force. The preloaded hinges therefore control the folding kinematics. This scheme has two primary disadvantages. One is the relatively stronger (and therefore heavier) boom required to overcome the hinge forces. The other is that in order for the hinge preload to be effective, panels must have sufficient rigidity to remain relatively flat between the hinges during deployment and restowage. Otherwise the panels themselves would bend and pivot away between hinges, thereby defeating the refolding attempt.
Previous attempts to utilize thin film panels have not been designed such that any stored energy in the hinges is propagated any significant distance in the panels from the hinges, primarily due to the fact that the thin film panels lack sufficient bending rigidity. The prior art has attempted to control thin film panel dynamics with intermediate connections between the extensible boom and/or linear guide tubes and wires and the panels. These methods result in considerable mechanical complexity, increased weight and potential low reliability of deployment and retraction due to the greater number of connections and sliding joints of this approach.
Another problem which has been inherent in the design of foldable solar panels for spacecraft use has been the requirement to place a "blanket" between the panels in the folded position in order to prevent face to face contact between the optical covers of the solar cells. This blanket adds weight to the panel assembly in the collapsed position and is a further reliability problem during deployment since it must be projected far enough from the panels during deployment to prevent its interfering with the solar panel function.